Solid propellant rocket engine control and missile configurations

ABSTRACT

The shaping of solid propellant charges for rocket engines to provide increased acceleration at missile launch without exceeding maximum acceleration near burnout while minimizing missile size and weight. Illustrative is a long-range ballistic missile having multiple rocket engine stages arranged in tandem, in which each stage has a reaction nozzle and associated propellant charge having a shape which is coordinated with the missile flight trajectory including a substantially conical or frusto-conical configuration providing a burning surface which decreases in area with burning time. Variable area reaction nozzles can be employed in each stage in combination with the conically-shaped propellant charges to optimize engine efficiency and increase engine versatility.

llnited States Patent Parilla Dec. 25, 1973 SOLID PROPELLANT ROCKETENGINE 2,206,809 7/1940 Denoix 1o2/49.7 x CONTROL AND MISSILE 2,489,95311/1949 Burney 60/242 X CONFIGURATIONS Inventor: Arthur R. Parilla, PO.Box 127,

Mountain Lakes, NJ. 07046 Filed: Dec. 16, 1971 Appl. No; 208,996

Related U.S. Application Data Continuation of Ser. No. 784,818, Dec. 18,1968, which is a division of Ser. No. 607,068, Jan. 3, 1967, Pat. No.3,489,373, which is a continuation-in-part of Ser. No. 302,222, June 14,1963, abandoned.

[52] U.S. Cl. 244/3.22, 60/234, 60/254, 102/34.5, l02/49.8, 102/100 [51]Int. Cl. F4lg 7/00, F42b 5/18 [58] Field of Search 244/321, 3.22;102/34.5, 37.7, 49.3, 49.7, 49.8, 99, 100, 101, 104; 60/234, 237, 254,256

[56] References Cited UNITED STATES PATENTS 3/1950 Chandler 60/256 XPrimary Examiner-Benjamin A. Borchelt Assistant Examiner.lames M. HanleyAttorneyl-larry C. Marcus 5 7] ABSTRACT The shaping of solid propellantcharges for rocket engines to provide increased acceleration at missilelaunch without exceeding maximum acceleration near burnout whileminimizing missile size and weight. 11- lustrative is a long-rangeballistic missile having multiple rocket engine stages arranged intandem, in which each stage has a reaction nozzle and associatedpropellant charge having a shape which is coordinated with the missileflight trajectory including a substantially conical or frusto-conicalconfiguration providing a burning surface which decreases in area withburning time. Variable area reaction nozzles can be employed in eachstage in combination with the conically-shaped propellant charges tooptimize engine efficiency and increase engine versatility.

2 Claims, 1 Drawing Figure PATENTEDHEBZS lava 3. 780.968

INVENTOR Arthur R.P0rill0 wgw wzw 46 750i ATTORNEYS.

SOLID PROPELLANT ROCKET ENGINE CONTROL AND MISSILE CONFIGURATIONSRELATED PATENTS AND APPLICATIONS This is a continuation, of applicationSer. No. 784,818, filed Dec. 18, 1968, and now abandoned.

This application is a division of applicant's prior co pendingapplication Ser. No. 607,068, filed Jan. 3,

1967 for Missile Configurations, Controls and Utilization Techniqueswhich is a continuation-in-part of applicants prior copendingapplication Ser. No. 302,222, filed June 14, 1963 and now abandoned, forAircraft Missiles, Missile Weapons System and Space Craft. Alsocopending as a division of Ser. No. 607,068 is application Ser. No.767,583, filed Oct. 3, 1968 for Missile Configurations, Controls andUtilization Techniques.

The abandoned application, Ser. No. 302,222, was a copending division ofapplication Ser. No. 701,571, filed Dec. 9, 1957 and now U.S. Pat. No.3,094,072, granted June 18, 1963. An additional application of theapplicant, Ser. No. 860,304, is also related to the instant application,being another continuation-in-part of Ser. No. 701,571, and now U.S.Pat. No. 3,151,446.

BACKGROUND AND SUMMARY OF THE INVENTION The shaping of solid propellantgrains as a means of controlling the rate of gas formation duringburning to obtain preselected thrust-time characteristics has been awell-known expedient in the design of rocket engines. In the most recentpast the emphasis in this area has been on the provision of constant orneutral burning propellant grains in which the rate of gas formationremains substantially uniform with burning time. Hence, the developmentof the so-called star and wagon wheel shaped grains.

While neutral burning grains are advantageous in some respects, thereare serious disadvantages inherent in their use. A neutral burninggrain, in providing a uniform rate of gas formation, results in aconstant or fixed thrust throughout the total burning time. It is wellknown that when missiles are equipped with constant thrust rocketengines, the missile acceleration increases as the missile massdecreases due to propellant consumption. Since the weight of thepropellant constitutes a very large portion of the total missile, alarge variation in missile acceleration can occur with constant thrustrocket engines.

Attempts have been made in the past to provide regressive burninggrains; i.e., grains whose inherent rate of gas formation decreases withburning time. Efforts in this connection have been concentrated on theformation of rod-shaped grains in which burning takes place on the outercylindrical surface. Such configurations have proven highly unsuccessfuldue in large measure to added weight resulting from the need for supportof the rod-shaped grain and insulation of the missile case walls. Infact, the mass ratios of such rockets have been so low as to precludepractical utilization.

Most rocket engines today are designed to operate at fixed thrustlevels. Among the objects of the instant invention are the following:

To provide a solid propellant rocket engine which provides increasedacceleration at missile launch while inherently limiting maximumacceleration near burnout.

To provide a rocket engine characterized by the above whose mass ratio,i.e., the ratio of the mass of the propellant to total mass, ismaximized.

To provide a rocket engine characterized by the above which makesmaximum utilization of the propellant for producing thrust during allphases of missile flight.

' To provide a rocket engine characterized by the above having athrust-time characteristic which is coordinated with the rate ofdecrease of the mass of the missile as the propellant is being consumed.

To provide a solid propellant rocket engine characterized by the abovehaving a thrust-time characteristic which is coordinated with themissile flight trajectory.

To provide a solid propellant rocket engine characterized by the abovehaving reduced overall size and weight.

To provide improved anti-ballistic missile weapon systems for localdefense against ballistic missile attack and for improved ground-to-airdefense.

The foregoing and other objects, features and advantages areaccomplished in accordance with the instant invention by the provisionof a solid propellant rocket engine whose grain is tapered or conical inconfiguration to provide a rate of gas formation which decreases withburning time. In the illustrative embodiment, a missile has a pluralityof rocket stages in tandem, each stage having a tapered or substantiallyconical or frusto-conical end burning propellant grain to provide aburning surface which decreases in area with burning time, the separategrains also conjointly defining an overall conical or frusto-conicalgeometrical configuration. It is preferred that the outer case of themissile also have a substantially conical shape conforming to theconfiguration of the propellant grain or grains.

In one preferred embodiment, each stage has a variable area reactionnozzle responsive to variations in combustion chamber pressure tomaintain such pressure at a constant while optimizing the nozzle thrustcoefficient as a function of missile altitude.

It can be seen that a self-throttling rocket engine of reduced weightand size is provided which automatically coordinates thrust with otherparameters to obtain maximum utilization of propellant fuel forincreased engine efficiency and flexibility while providing increasedacceleration at launch and reduced acceleration near burnout.

Reference will now be had to the drawings in which the single FIGURE isa partially diagrammatic elevational view of a long-range ballisticmissile embodying the invention, with portions of the missile beingbroken away and shown in section.

In the missile, which provides improved perfonnance, 251 is the warhead,and 252 is a guidance system supported by the airframe 253. The secondstage rocket 254 contains a solid propellant, 255, within a rocket case,256. Additional propellant 257 may provide higher loading density,filling the space within the nozzle plug, 273. The variable area cowland expansion cone, 258 is attached by the bellows 271 and actuators 259to the case 256. Control system 260 for the missile may be mounted inthe second stage between the nozz-le 258 and the case, 256.

The first stage 261 and the second stage 254 are joined by the splitcollar 268 and the upper and lower clamps 269 and 270.

The first stage rocket contains propellant 262 within the case 263, withadditional propellant 264 filling the space within the nozzle plug 265,increasing loading density. The cowl and expansion cone 266 is attachedto the case by the bellows 267 and actuators 268, the details being aspreviously described.

The propellant grains for both first and second stages are designed toprovide the desired regressive characteristics. This, then, permitsautomatic reduction of thrust as a function of time, permitting higheracceleration at launch without exceeding maximum acceleration nearburnout. The grain regressivity is selected to permit automaticoperation of the variable nozzle area expansion ratio and, hence,improve nozzle thrust coefficient as a function of altitude, using timeas a common parameter. Both of these purposes may now be accomplishedwith no additional controls other than the means described forpositioning the nozzle cowl in the aforesaid copending application Ser.No. 607,068, such as the mechanical or fluid springs, or thecontrollable solid propellant systems described therein. Any of thesemethods, of course, will still provide minimum case weight by acting asconstant pressure devices independent of ambient temperatures; they willalso act to minimize thrust variation with ambient temperature changes,as described.

While the propellant grains may have any configuration so long as thedesired regressivity is obtained, for the purpose of illustration, anend-burning grain is shown which has a tapered or conical cross-sectionwhich reduces in area as burning progresses. In the illustration shownin the FIGURE, ignition may occur on both surfaces of the two grains 262and 264 is Stage 1 rocket engine. The burning surface then remainssubstantially constant for the first portion of the burning time, theprogressivity of grain 262 after ignition being reduced by theregressivity of grain 264. Small variation in grain geometryparticularly increased burning area, will have little serious effectsince the nozzle cowl will extend automatically, increasing throat area,eliminating pressure peaks. This permits greater freedom in attaininghigh loading densities.

When the first stage is ignited (by conventional means not shown) thenozzle cowl 266 assumes its normal design position, as by compression ofthe springs or actuators 268. In this position, the area expansion ratiomay be approximately 5, corresponding to optimum thrust coefficient atsea-level with a chamber pressure of 500 psia. As the missile gainsaltitude, the chamber pressure and mass flow rate remain constant duringan initial period, as long as burning surface remains constant. After atime, t,, the missile would be at an altitude at which the pressureratio will have increased to a value causing the jet diameter toincrease to the full nozzle diameter. The area ratio, with constant cowlposition, may now be 10, corresponding to an optimum nozzle coefficientat an altitude of 25,000 feet. As the missile continues to climb, thegrain becomes regressive. The reduced value of K (the ratio of burningsurface to nozzle throat area) as burning surface reduces would tend toreduce chamber pressure. The springs (or control system) 268 then causethe nozzle cowl 266 to retract, the reduced throat area reducing thrustas constant pressure is maintained, and increasing the expansion ratio.As an example, when burning has progressed to a section where thepropellant burning surface was approximately one-half of its initialvalue, the

mass flow rate is reduced to one-half its initial value; the throat areaalso reduces to one-half of its initial value thereby again doubling thenozzle expansion ratio approaching a value of 20. This would nowcorrespond to an optimum ratio and optimum thrust coefficient atapproximately 45,000 feet altitude with the constant chamber pressure of500 psia.

Referring to standard references showing the relationship of nozzlethrust coefficient as functions of pressure ratio and area ratio, it maybe seen that the variable area expansion ratio increases the nozzlethrust coefficient and, hence, the propellant specific impulse by morethan 10 percent at upper altitudes, while still maintaining optimumnozzle coefficient at sea-level. In view of the large propellant massesrequired for long range missiles, this permits substantial reduction inactual propellant weight, with corresponding reduction in missile sizeand case weight.

When this same feature is used in upper stages, together with maximumnozzle expansion ratios, the higher specific impulse makes possiblereduced propellant weight of upper stages. This saving is magnified byreduced weight of lower stages, substantially reducing take-off weight,and size of the missile as a whole.

Thus, it may be seen that by the use of a regressive burning propellant,whose design is coordinated with the missile flight trajectory, simplecontrol means may be provided whereby the rocket engine automaticallyincreases the nozzle expansion ratio and thrust coefficient as afunction of altitude. The resulting grain regressivity then also reducesthrust automatically as propellant is consumed, inherently providingthrottling to restrict maximum acceleration as the mass of the missilereduces.

Upon first stage burnout, the clamps 269 and 270 are released. This maybe accomplished by explosive bolts, or by a quick release mechanismresponsive to loss of chamber pressure in the first stage. The ring 268,which is split circumferentially into two segments, falls away providingclearance for the large expansion cone of the second stage nozzle 258 tooscillate for vector control.

The second stage rocket engine is made regressive to accomplish the samevariable expansion ratio and also to restrict maximum acceleration, aswith the first stage. The @QQI'ILIQQ ratiomay increase from, say, 40

after ignition, to twice that value as the r nissile continues to climb.

The second stage nozzle is designed to utilize the maximumcross-sectional area available within the missile envelope, thusproviding maximum nozzle thrust coefficient at the extremely highpressure ratios in and beyond the upper atmosphere. It may be of theinternal expansion type, described and illustrated in FIG. 4 of theaforesaid copending application Ser. No. 607,068. Since second stageignition occurs only after high altitudes are reached, the high pressureratios result in high Prandtl-Meyer angles. The nozzle may thereforehave a large angle at the throat, followed by a contoured divergentsection.

The flow undergoes expansion beyond the throat, followed by somerecompression as it is directed aft in an axial direction. Thus,extremely large area expansion ratios may be obtained within relativelyshort nozzle lengths.

Overall missile length may be reduced by nesting the upper head closureof the first stage within the nozzle expansion cone of the second stage.Clearance for the plug tip of the second stage nozzle may be provided bya recess in the first stage head closure, if required, as at 272, andfor alignment of successive stages.

The double release clamps 269 and 270 and split ring 268 may be replacedby a single clamp or equivalent fitting. This would reduce the nozzleexpansion cone to insure sufficient clearance on separation. The highernozzle thrust coefficient and higher specific impulse with the largerexpansion cone would save propellant weight when the two clamps areused.

Thrust vector control, as well as thrust termination may, of course, beincorporated with the variable area nozzles as previously described inthe aforesaid copending application Ser. No. 607,068.

Multiple stages exceeding two may, of course, be used. The conical grainis illustrative only of regressive burning grain. A cylindrical missilehaving cylindrical grains may, of course, be used, the desiredregressivity being obtained by other grain geometries.

The propellant burning rate may be made regressive, rather than burningsurface. This may be controlled by varying propellant composition, useof one or more grains; or by controlling or varying the length of wiressometimes embedded within the propellant grain to increase the burningrate; the number of wires per unit of cross-section being varied alongthe length of a cylindrical grain.

As an aid to understanding the invention, the applicants prior copendingand related patents and applications previously cited herein should bereviewed. These patents and applications illustrate instrumentalitiesand techniques for varying the throat area and other geometricparameters of nozzles employed in thrust producing devices includingexamples of a number of embodiments for varying the throat area andstream orientation in rocket engine nozzles to achieve variations inexpansion ratio, chamber pressure, pressure ratio, propellant specificimpulse, thrust magnitude and direction, and related parameters.

The invention in its broader aspects is not limited to the steps,arrangements, instrumentalities and combinations shown and describedherein for illustrative purposes, and departures may be made therefromby those skilled in the art without departing from the spirit and scopeof the invention as defined in the appended claims and withoutsacrificing its chief advantages.

What is claimed is:

l. A missile comprising an outer case defining a combustion chamber fora propellant grain; a regressivebuming propellant grain in saidcombustion chamber; and, a variable throat area reaction noule movablyconnected to said case for receiving gas issuing from said combustionchamber, said nozzle being movable in response to the mass flow rate ofgas issuing from said chamber to vary said throat area for maintaining asubstantially constant pressure within said combustion chamber fromSea-Level to an attained altitude as the rate of gas formation in, andthe mass flow rate of gas issuing from, said chamber decreases withtime, the regressive-buming characteristics of said grain being selectedto maintain a substantially constant ratio between the mass flow rate ofgas issuing from said chamber and the throat area of said nozzle, withtime, and to provide sufficient initial generation of propellant gasesto accelerate said missile during launch without exceeding maximumacceleration near burnout, whereby, at constant chamber pressure fromSea-Level to the attained altitude, the thrust coefficient of saidnozzle increases from an optimum coefficient at Sea- Level to an optimumcoefficient at the attained altitude.

2. A missile comprising an outer case, a plurality of separable rocketstages arranged in tandem for sequential firing in propelling saidmissile, each of said stages having a reaction nozzle in operativecommunication stage.

1. A missile comprising an outer case defining a combustion chamber fora propellant grain; a regressive-burning propellant grain in saidcombustion chamber; and, a variable throat area reaction nozzle movablyconnected to said case for receiving gas issuing from said combustionchamber, said nozzle being movable in response to the mass flow rate ofgas issuing from said chamber to vary said throat area for maintaining asubstantially constant pressure within said combustion chamber fromSea-Level to an attained altitude as the rate of gas formation in, andthe mass flow rate of gas issuing from, said chamber decreases withtime, the regressive-burning characteristics of said grain beingselected to maintain a substantially constant ratio between the massflow rate of gas issuing from said chamber and the throat area of saidnozzle, with time, and to provide sufficient initial generation ofpropellant gases to accelerate said missile during launch withoutexceeding maximum acceleration near burnout, whereby, at constantchamber pressure from Sea-Level to the attained altitude, the thrustcoefficient of said nozzle increases from an optimum coefficient atSea-Level to an optimum coefficient at the attained altitude.
 2. Amissile comprising an outer case, a plurality of separable rocket stagesarranged in tandem for sequential firing in propelling said missile,each of said stages having a reaction nozzle in operative communicationwith an associated solid propellant charge within said outer case, andsaid propellant charges being formed to individually and conjointly havea generally conical geometrical configuration to provide a burningsurface which decreases in area with the burning time of each stage.